1. Technical Field
This invention relates to airfoils for gas turbines in general, and to hollow airfoils having apparatus for cooling the leading edge and establishing film cooling along the surface of the airfoil in particular.
2. Background Information
In the turbine section of a gas turbine engine, core gas travels through a plurality of stator vane and rotor blade stages. Each stator vane or rotor blade has an airfoil with one or more internal cavities surrounded by an external wall. The suction and pressure side portions of the external wall extend between the leading and trailing edges of the airfoil. Stator vane airfoils extend spanwise between inner and outer platforms and rotor blade airfoils extend spanwise between a platform and a blade tip.
High temperature core gas (which includes air and combustion products) encountering the leading edge of an airfoil will diverge around the suction and pressure side portions of the airfoil, with some of the gas impinging on the leading edge. The point along the airfoil where the velocity of the core gas flow decelerates to zero (i.e., the impingement point) is referred to as the stagnation point. There is a stagnation point at every spanwise position along the leading edge, and collectively those points are referred to as the stagnation line. Air impinging on or adjacent the leading edge is subsequently diverted around either side of the airfoil. The precise location of each stagnation point along the leading edge is a function of the angle of incidence of the core gas relative to the chordline of the airfoil, for both rotor and stator airfoils. In addition to the angle of incidence, the stagnation point of a rotor airfoil is also a function of the rotational velocity of the airfoil and the velocity of the core gas. Given the curvature of the leading edge, the approaching core gas direction and velocity, and the rotational speed of the airfoil (if any), the location of the stagnation points along the leading edge can be readily determined by means well-known in the art. In actual practice, rotor speeds and core gas velocities vary depending upon engine operating conditions as a function of time and position along the leading edge. As a result, some movement of the stagnation points (or collectively the stagnation line) along the leading edge can be expected during operation of the airfoil.
Cooling air, typically extracted from the compressor at a temperature lower and pressure higher than the core gas passing through the turbine section, is used to cool the airfoils. The cooler compressor air provides the medium for heat transfer and the difference in pressure provides the energy required to pass the cooling air through the stator or rotor stage.
In many cases, it is desirable to establish a film of cooling air along the surface of the stator or rotor airfoil by bleeding cooling air out of cooling holes. The term "bleeding" reflects the small difference in pressure motivating the cooling air out of the internal cavity of the airfoil. The film of cooling air traveling along the surface of the airfoil directs the flow of high thermal energy hot gas away from the airfoil, increases the uniformity of the cooling, and thermally insulates the airfoil from the passing hot core gas. A person of skill in the art will recognize, however, that film cooling is difficult to establish and maintain in the turbulent environment of a gas turbine.
A known method of establishing film cooling involves positioning cooling holes in or adjacent the leading edge of an airfoil in a "showerhead" arrangement. The showerhead typically includes a row of cooling holes on either side of the leading edge. The cooling holes are angled aft and are often diffused to facilitate film formation. In some cases, the showerhead includes a row of holes positioned directly on the leading edge. U.S. Pat. No. 5,374,162 discloses an example of such an arrangement.
One problem associated with using holes to create a cooling air film is the film's sensitivity to pressure difference across the holes. Too great a pressure difference across a cooling hole will cause the air to jet out into the passing core gas rather than aid in film formation. Too small a pressure difference will result in negligible cooling air flow through the hole, or worse, an in-flow of hot core gas. Both cases adversely affect film cooling effectiveness. Another problem associated with using holes to establish film cooling is that cooling air is dispensed from discrete points along the span of the airfoil, rather than along a continuous line. The gaps between cooling holes, and areas immediately downstream of those gaps, are exposed to less cooling air than are the holes and the spaces immediately downstream of the holes, and are therefore more susceptible to thermal distress. Another problem associated with using holes to establish film cooling is the stress concentrations that accompany each hole. Stress concentrations develop when loads (typically resulting from dynamic forces or thermal expansion) are carried by narrow expanses of material extending between adjacent holes. Film cooling effectiveness generally increases when the cooling holes are closely packed and skewed aft at a shallow angle relative to the external surface of the airfoil. Skewed, closely packed apertures, however, are more prone to stress concentrations.
Some prior art configurations have cooling holes disposed in the leading edge aligned with an average stagnation line, that extend perpendicular to the external surface of the airfoil. Such a cooling hole arrangement can experience an asymmetrical cooling air distribution. For example, an actual stagnation line shift to one side of a row of cooling holes can urge exiting cooling air to one side of the row, consequently leaving the opposite side starved of cooling air. The fact that the stagnation line can and does shift during airfoil operation illustrates that locating holes on the average stagnation line will not remedy all cooling air distribution problems. Cooling holes extending perpendicular to the external surface and skewed spanwise do not resolve the potential for asymmetrical cooling air distribution.
What is needed is an apparatus that provides adequate cooling along the leading edge of an airfoil, one that accommodates a variable position stagnation line, and one that promotes a uniform and durable cooling air film downstream of the leading edge on both sides of the airfoil.